1. Field of the Invention
The present invention relates to a rocket propulsion system which includes an array of eccentric rocket engines that function as thrusters and/or divert engines for controlling the direction of flight of the rocket assembly.
2. Description of the Related Art
In the art relating to aerospace vehicles, various types of propulsion systems are known. A large number of these systems are constructed to provide thrust as well as attitude control. Each of these systems has its own characteristic advantages and disadvantages.
However, existing propulsion systems and attitude and control systems have certain drawbacks in that they are unable to achieve a combination of a high energy density and controllable thrust found with solid propulsion system and, at the same time, exhibit the reliable shut down and re-ignition capabilities of a liquid or hybrid system.
It is an object of the present invention to provide a rocket motor assembly including a propulsion system having a high propulsion capacity, yet which allows individual engines or sets of engines to be independently shut down and reliably re-ignited as desired to achieve accurate attitude control.
The propulsion systems according to this invention include an array of eccentric rocket engines each having a combustion chamber. As referred to herein, eccentric means that the rocket engines are offset from the longitudinal axis of the rocket motor assembly so that firing of selected ones or groups of the eccentric rocket engines permits attitude control over the rocket assembly. (Although the eccentric rocket engines are individually offset from the longitudinal axis of the rocket assembly, the eccentric motors can, for example, collectively form a concentric ring about the longitudinal axis of the rocket assembly, so that the simultaneous firing of certain or all of the eccentric rocket motors thrusts the rocket assembly without attitude adjustment.) As referred to herein, attitude control means influencing the pitch, yaw, and/or roll of the rocket assembly in flight.
The propulsion systems further include one or more oxidizer-fluid sources, e.g., storage tanks, in operative communication with the eccentric rocket engines to permit oxidizer fluid to be supplied from the oxidizer-fluid sources to the combustion chambers of the eccentric rocket engines. For example, one oxidizer-fluid source can be provided that delivers oxidizer fluid to each of the eccentric rocket engines. Alternatively, a plurality of oxidizer-fluid sources can each communicate with a corresponding one or groups of eccentric rocket engines. Representative oxidizer fluids include solutions, slurries, or gels containing primary oxidizers such as hydrogen peroxide, nitrogen tetroxide, inhibited red fuming nitric acid (IRFNA), hydroxyl ammonium nitrate (HAN), ammonium nitrate (AN), ammonium perchlorate, hydroxyl ammonium perchlorate (HAP), and other oxidizers well known in the art.
One or more ignition-fluid sources, e.g., storage tanks, are placed in operative communication with the eccentric rocket engines to permit ignition fluid to be delivered to the combustion chambers of the eccentric rocket engines. Again, one ignition-fluid source can deliver ignition fluid to each of the eccentric rocket engines. Alternatively, a plurality of ignition-fluid sources can each deliver ignition fluid to a corresponding one or group of eccentric rocket engines.
In a first preferred embodiment of the invention, the eccentric rocket engines are hybrid rocket engines housing solid fuel grains. In this embodiment, when hot ignition fluid, typically in a gaseous state, is introduced into the combustion chambers together with the oxidizer fluid, a combustion reaction ensues. Once ignition has occurred, the flow of hot ignition fluid into the combustion chambers can be terminated without stopping the combustion reactions, since a solid fuel source is already present in the combustion chamber. The combustion reaction can be terminated by terminating the flow of oxidizer fluid from the oxidizer-fluid source to the combustion chamber of the hybrid rocket engines. The solid fuel grain can be devoid of solid oxidizer or can contain small amounts of solid oxidizer, so long as the solid oxidizer is not present in the grain in sufficient concentrations to create a self-deflagrating reaction.
In a second preferred embodiment of the invention, the eccentric rocket engines are bi-fluid rocket engines. In this embodiment, the ignition/fuel source, which may be, by way of example, a gas generator containing a solid propellant that is converted by controlled combustion into a hot gas, can deliver the fuel component for the combustion reaction and simultaneously supply the heat necessary for ignition. In the bi-fluid embodiment, the ignition/fuel source generally should be supplied to the combustion chamber at a higher flow rate than for the hybrid rocket engine of the first preferred embodiment, since unlike a hybrid rocket engine, a bi-fluid engine does not contain a solid fuel grain in the combustion chamber. When the ignition/fuel fluid and oxidizer fluid are both delivered into the combustion chambers, a combustion reaction ensues. The combustion reaction can be stopped by terminating the flow of oxidizer fluid and/or ignition/fuel fluid to the combustion chamber of the bi-liquid rocket engines.
In accordance with a third embodiment of the invention, the eccentric rocket engines are single-fluid engines, in which the fluid supplied to all of the engines is generated and delivered from a single or multiple sources of combustion products, with at least some of said sources being connected to at least two eccentric rocket engines. In this embodiment, firing of the eccentric rocket engines is controlled by permitting and terminating the flow of combustion products to the eccentric rocket engines individually.
It is also within the scope of this invention to use combinations of hybrid rocket engines and bi-fluid rocket engines. Preferably, the eccentric rocket engines are connected to the oxidizer-fluid source and ignition-fluid source with respective control valves, which more preferably permit the variable throttling, shut down, and re-starting of selected ones or groups of the eccentric rocket engines independently of one another.
The propulsion system of this invention optionally, and in some embodiments preferentially, includes an axial primary engine or a plurality of primary engines constructed and arranged to collectively produce thrust force along the axial or longitudinal direction. When present, the primary engine or engines are preferably the main propulsion source so as to produce a higher level of thrust than the eccentric engines in the array produce individually. The primary engine or engines can be hybrid engine(s), reverse hybrid engine(s), bi-fluid engine(s), self-deflagrating solid propellant engine(s), or dual chamber solid engine(s). The primary engine or engines can receive the oxidizer and/or fuel from the oxidizer-fluid source(s) and fuel-fluid source(s), respectively.
In accordance with a preferred modification to the invention, the propulsion system further includes at least one cooling device operatively connecting the ignition-fluid source(s) to the oxidizer-fluid source(s). A portion of the ignition fluid from the ignition-fluid source is sent through the cooling device to lower the temperature of the ignition fluid, and then is used to pressurize fluid oxidizer in the oxidizer-fluid source. Although various mechanical and pneumatic set ups can be envisioned, by way of example, the oxidizer source can be provided with a piston or an expandable or collapsible bladder or like device for keeping the oxidizer liquid and cooled ignition fluid separate.
There is a possibility, in accordance with a less preferred embodiment of the present invention, of a separate cold and warm gas oxidizer pressurization system for the oxidizer-fluid source by means within the purview of a skilled artisan. While this option offers a degree of simplicity and reliability, it lacks the improved packaging and performance of the preferred embodiment.
There is also a further possibility, in accordance with a less preferred embodiment of the present invention, of not cooling the propellant gas which is delivered to pressurize the oxidizer-fluid source. In this case, the hot gas would decompose a small amount of oxidizer in the oxidizer-fluid source and the decomposed oxidizer would then pressurize the oxidizer-fluid source. While this option offers certain benefits, it also requires careful control of the amount of decomposition achieved in order to pressurize the oxidizer-fluid source.
Systems according to the invention provide improved packaging and performance over known systems, such as hydrazine-based or solid propellant-based propulsion/attitude control systems. The use of hybrid and/or bi-fluid propellant technology provides reliable re-ignition and throttling capabilities compared to pure solid fuel engines, while also being extremely safe and virtually explosion proof. The igniter consists of combusted propellant, which functions either as the ignition fluid for hybrid technology or as both the ignition fluid and fuel source for bi-fluid technology.
Other objects, aspects and advantages of the invention will be apparent to those skilled in the art upon reading the specification and appended claims which, when read in conjunction with the accompanying drawings, explain the principles of this invention.